Shape correcting components

The present invention relates to correcting and setting the shape of cast components. It is particularly, though not exclusively, concerned with correcting the shape of brittle, expensive components that require a high level of precision.

A gas turbine engine, as shown in <figref idrefs="DRAWINGS">FIG. 1</figref>, comprises an air intake <b>12</b> and a propulsive fan <b>14</b> that generates two airflows A and B. The gas turbine engine <b>10</b> comprises, in axial flow A, an intermediate pressure compressor <b>16</b>, a high pressure compressor <b>18</b>, a combustor <b>20</b>, a high pressure turbine <b>22</b>, an intermediate pressure turbine <b>24</b>, a low pressure turbine <b>26</b> and an exhaust nozzle <b>28</b>. A nacelle <b>30</b> surrounds the gas turbine engine <b>10</b> and defines, in axial flow B, a bypass duct <b>32</b>.

Typically the blades of the low pressure turbine <b>26</b> are cast from nickel alloys. Casting does not always result in a perfectly formed component and thus some correction of the shape is required. This can be performed relatively easily and cheaply by mechanical plastic deformation. However, there is a requirement to replace nickel alloys with intermetallics such as titanium aluminide alloys to reduce the weight of the low pressure turbine without compromising the strength of the blades. Gamma titanium aluminide (γ-TiAl) is a desirable alloy for low pressure turbine blades. However, it is a relatively brittle material and therefore cannot be deformed using mechanical plastic deformation.

One disadvantage of this alloy is that low pressure turbine blades cast from intermetallics such as γ-TiAl must either suffer very low yield, due to the blade being imperfectly shaped, or must be cast oversize and then machined to the desired shape. In either case this is expensive, time consuming and wasteful.

The present invention seeks to provide a method of forming a perfectly shaped component that seeks to address the aforementioned problems.

Accordingly the present invention provides a method of forming a component comprising the steps of casting a component; placing the component adjacent a mould surface; and creep deforming the component during the application of heat and pressure to conform at least a part thereof to the mould surface.

Preferably the component comprises titanium aluminide alloy, forms of silicide based on niobium or molybdenum, or ceramic.

Preferably the applied pressure comprises isostatic pressure. More preferably the hot isostatic pressure is applied via a secondary particulate material.

Preferably the component and mould surface are wrapped in a foil to prevent infiltration between the component and mould surface by the secondary particulate material. More preferably the foil is yttria coated.

Preferably the component is a turbine blade for a gas turbine engine. More preferably the component is a low pressure turbine blade.

Preferably the creep mould is ceramic. More preferably the creep mould comprises yttria face coated alumina or silica.

Preferably the component is cast in a net-shape mould.

The present invention will be more fully described by way of example with reference to the accompanying drawings, in which:

<figref idrefs="DRAWINGS">FIG. 1</figref> is a sectional side view of a gas turbine engine.

<figref idrefs="DRAWINGS">FIG. 2</figref> and <figref idrefs="DRAWINGS">FIG. 3</figref> are schematic side views of a component before and after creep deformation according to a first embodiment of the present invention.

<figref idrefs="DRAWINGS">FIG. 4</figref> and <figref idrefs="DRAWINGS">FIG. 5</figref> are schematic side views of a component before and after creep deformation according to a second embodiment of the present invention.

The method of the present invention is described with reference to the first embodiment shown in <figref idrefs="DRAWINGS">FIG. 2</figref> and <figref idrefs="DRAWINGS">FIG. 3</figref>. In a first step of the method, a turbine blade <b>34</b> is cast from γ-TiAl in a net-shape mould. This results in a blade <b>34</b> that is close to the desired shape and usually contains internal gas and shrinkage porosity.

In a second step of the method of the present invention, the turbine blade <b>34</b>, having a pressure surface <b>36</b> and a suction surface <b>38</b>, is placed onto a creep mould <b>40</b> having a mould surface <b>42</b> that defines the desired shape of the pressure surface <b>36</b> of the turbine blade <b>34</b>. The pressure surface <b>36</b> of the blade <b>34</b> is placed against, but does not exactly conform to, the mould surface <b>42</b> of the creep mould <b>40</b>, as is shown in <figref idrefs="DRAWINGS">FIG. 2</figref>. It does not exactly conform due to the imperfect shape of the cast turbine blade <b>34</b>. The turbine blade <b>34</b> and the creep mould <b>40</b> are preferably wrapped in an inert foil <b>43</b>, such as mild steel foil. To avoid contamination of the blade <b>34</b> by the foil <b>43</b>, a releasing agent such as a thin yttria coating may be used.

The creep mould <b>40</b> is preferably a ceramic component so that it is unaffected by the heat supplied thereto in a later step of the method. Preferably, the ceramic is yttria face coated alumina or silica, which is similar to the material used to form casting moulds.

In a third step of the method, the arrangement <b>44</b>, comprising the turbine blade <b>34</b>, the creep mould <b>40</b> and the foil <b>43</b>, is placed inside a canister comprising a deformable wall inside a hot isostatic pressure (HIP) chamber. Between the deformable wall and the arrangement <b>44</b> inside is a solid particulate material that can transfer heat and isostatic pressure from supply means located externally of the deformable wall to the arrangement <b>44</b>. The foil wrapping <b>43</b> prevents the solid particulate material infiltrating the gap between the turbine blade <b>34</b> and the creep mould <b>40</b>. Typically the isostatic pressure is applied to the deformable wall by argon gas impingement, which can also be heated, but other known methods can be used with equal felicity.

In a fourth step of the method, heat and isostatic pressure are applied to the canister inside the HIP chamber to consolidate the component and close all porosities. During the application of heat and isostatic pressure in the HIP chamber the turbine blade <b>34</b> also deforms through the mechanism of creep. Since the creep mould <b>40</b> retains its shape throughout the HIP process, the turbine blade <b>34</b> deforms under its own weight so that its pressure surface <b>36</b> conforms to the shape of the mould surface <b>42</b> of the creep mould <b>40</b> as shown in <figref idrefs="DRAWINGS">FIG. 3</figref>. The arrangement <b>44</b> is then removed from the HIP chamber and the canister and the turbine blade <b>34</b> can be removed from the creep mould <b>40</b>. Typically the turbine blade <b>34</b> requires further processing, for example the addition of cooling holes, as is well known in the art.

Hence, this method provides a turbine blade <b>34</b> that is fully consolidated and has the desired shape. This substantially reduces or eliminates the waste associated with a need to scrap imperfectly shaped blades <b>34</b> or to cast an oversize component and machine away waste material until the desired shape and size is obtained.

A typical turbine blade <b>34</b> is around 400 mm long, indicated by arrows y in <figref idrefs="DRAWINGS">FIG. 2</figref>. The furthest distance between the pressure surface <b>36</b> of the turbine blade <b>34</b> and the mould surface <b>42</b> of the creep mould <b>40</b> is typically a few millimetres, up to around 10 mm and indicated by arrows x. Thus the method of the present invention is able to correct the shape of a cast γ-TiAl turbine blade <b>34</b> by around 10 mm during the HIP step.

A second embodiment of the present invention is shown in <figref idrefs="DRAWINGS">FIG. 4</figref> and <figref idrefs="DRAWINGS">FIG. 5</figref> in which like reference numerals are used for like components. As in the first embodiment described above, a turbine blade <b>34</b> having pressure <b>36</b> and suction <b>38</b> surfaces is placed onto a first creep mould <b>40</b> so that the pressure surface <b>36</b> of the blade <b>34</b> is adjacent to the mould surface <b>42</b> of the first creep mould <b>40</b>. A second creep mould <b>46</b>, having a mould surface <b>48</b> defining the desired shape of the suction surface <b>38</b> of the turbine blade <b>34</b>, is placed onto the turbine blade <b>34</b> so that the mould surface <b>48</b> thereof is adjacent to the suction surface <b>38</b> of the turbine blade <b>34</b>. As with the first embodiment, the turbine blade <b>34</b> and creep moulds <b>40</b>, <b>46</b> are preferably wrapped in an inert foil <b>43</b> such as mild steel foil.

The arrangement <b>50</b>, comprising the turbine blade <b>34</b> and the first and second creep moulds <b>40</b>, <b>46</b>, is placed inside a canister within a HIP chamber as described with respect to the first embodiment. Heat and isostatic pressure are applied to the canister inside the HIP chamber, preferably by heated argon gas, to consolidate the blade <b>34</b> and to close the porosities. The combined weight of the blade <b>34</b> and the second (upper) creep mould <b>46</b> also causes the turbine blade <b>34</b> to creep. The first and second creep moulds <b>40</b>, <b>46</b> constrain the blade <b>34</b> to deform during creep to conform to the shape of the adjacent creep mould. Hence, the turbine blade <b>34</b> creeps to the shape shown in <figref idrefs="DRAWINGS">FIG. 5</figref>.

The arrangement <b>50</b> can be removed from the canister and the HIP chamber and the turbine blade <b>34</b> extracted from between the creep moulds <b>40</b>, <b>46</b>. Further processing may be required as discussed in relation to the first embodiment.

Although the method of the present invention has been described with respect to the shape correction and setting of a turbine blade <b>34</b>, it may be applied to other components of a gas turbine engine, for example low pressure turbine stators and high pressure compressor stators and blades.

Although the canister has been described with a deformable wall surrounding a solid particulate material for transferring the heat and isostatic pressure, other known methods of HIP treating a component could be employed. For example, direct application of heat and isostatic pressure to a sealed foil assembly, although this has been found to be less efficacious than the indirect method described above.

Although the isostatic pressure applied to the deformable wall is described as via argon gas impingement, which can also be heated, other known methods can be used with equal felicity.

Although creep setting of intermetallics such as γ-TiAl has been described, the method of the present invention can also be used with other brittle materials such as ceramics and forms of silicide based on niobium or molybdenum. Such materials could be used for components in hotter parts of a gas turbine engine or in other applications.